In this paper,we consider a 2-degrees-of-freedom(DOF)helicopter system subject to long input delays and uncertain system parameters.To address the challenges including unknown system parameters and input delays in con...In this paper,we consider a 2-degrees-of-freedom(DOF)helicopter system subject to long input delays and uncertain system parameters.To address the challenges including unknown system parameters and input delays in control design,we develop an adaptive predictor-feedback control law to achieve trajectory tracking.Stability of the closed-loop system is further established,where the tracking errors are shown to converge towards zero.Through simulation and experiments on the helicopter system,we illustrate that tracking of a desired trajectory is achieved with the proposed control scheme.展开更多
Helicopter systems present numerous benefits over fixed-wing aircraft in several fields of application.Developing control schemes for improving the tracking accuracy of such systems is crucial.This paper proposes a ne...Helicopter systems present numerous benefits over fixed-wing aircraft in several fields of application.Developing control schemes for improving the tracking accuracy of such systems is crucial.This paper proposes a neural-network(NN)-based adaptive finite-time control for a two-degree-of-freedom helicopter system.In particular,a radial basis function NN is adopted to solve uncertainty in the helicopter system.Furthermore,an event-triggering mechanism(ETM)with a switching threshold is proposed to alleviate the communication burden on the system.By proposing an adaptive parameter,a bounded estimation,and a smooth function approach,the effect of network measurement errors is effectively compensated for while simultaneously avoiding the Zeno phenomenon.Additionally,the developed adaptive finite-time control technique based on an NN guarantees finitetime convergence of the tracking error,thus enhancing the control accuracy of the system.In addition,the Lyapunov direct method demonstrates that the closed-loop system is semiglobally finite-time stable.Finally,simulation and experimental results show the effectiveness of the control strategy.展开更多
The so-called coaxial compound helicopter features two rigid coaxial rotors,and possesses high-speed capabilities.Nevertheless,the small separation of the coaxial rotors causes severe aerodynamic interactions,which re...The so-called coaxial compound helicopter features two rigid coaxial rotors,and possesses high-speed capabilities.Nevertheless,the small separation of the coaxial rotors causes severe aerodynamic interactions,which require careful analysis.In the present work,the aerodynamic interaction between the various helicopter components is investigated by means of a numerical method considering both hover and forward flight conditions.While a sliding mesh method is used to deal with the rotating coaxial rotors,the Reynolds-Averaged Navier-Stokes(RANS)equations are solved for the flow field.The Caradonna&Tung(CT)rotor and Harrington-2 coaxial rotor are considered to validate the numerical method.The results show that the aerodynamic interaction of the two rigid coaxial rotors significantly influences hover’s induced velocity and pressure distribution.In addition,the average thrust of an isolated coaxial rotor is smaller than that of the corresponding isolated single rotor.Compared with the isolated coaxial rotor,the existence of the fuselage results in an increment in the thrust of the rotors.Furthermore,these interactions between the components of the considered coaxial compound helicopter decay with an increase in the advance ratio.展开更多
In this paper, a disturbance observer-based safe tracking control scheme is proposed for a medium-scale unmanned helicopter with rotor flapping dynamics in the presence of partial state constraints and unknown externa...In this paper, a disturbance observer-based safe tracking control scheme is proposed for a medium-scale unmanned helicopter with rotor flapping dynamics in the presence of partial state constraints and unknown external disturbances. A safety protection algorithm is proposed to keep the constrained states within the given safe-set. A second-order disturbance observer technique is utilized to estimate the external disturbances. It is shown that the desired tracking performance of the controlled unmanned helicopter can be achieved with the application of the backstepping approach, dynamic surface control technique, and Lyapunov method. Finally, the availability of the proposed control scheme has been shown by simulation results.展开更多
High-speed rotor rotation under the low-density condition creates a special low-Reynolds compressible flow around the rotor blade airfoil where the compressibility effect on the laminar separated shear layer occurs. H...High-speed rotor rotation under the low-density condition creates a special low-Reynolds compressible flow around the rotor blade airfoil where the compressibility effect on the laminar separated shear layer occurs. However, the compressibility effect and shock wave generation associated with the increase in the Mach number (M) and the trend change due to their interference have not been clarified. The purpose is to clear the compressibility effect and its impact of shock wave generation on the flow field and aerodynamics. Therefore, we perform a two-dimensional unsteady calculation by Computational fluid dynamics (CFD) analysis using the CLF5605 airfoil used in the Mars helicopter Ingenuity, which succeeded in its first flight on Mars. The calculation conditions are set to the Reynolds number (Re) at 75% rotor span in hovering (Re = 15,400), and the Mach number was varied from incompressible (M = 0.2) to transonic (M = 1.2). The compressible fluid dynamics solver FaSTAR developed by the Japan aerospace exploration agency (JAXA) is used, and calculations are performed under multiple conditions in which the Mach number and angle of attack (α) are swept. The results show that a flow field is similar to that in the Earth’s atmosphere above M = 1.0, such as bow shock at the leading edge, whereas multiple λ-type shock waves are observed over the separated shear layer above α = 3° at M = 0.80. However, no significant difference is found in the C<sub>p</sub> distribution around the airfoil between M = 0.6 and M = 0.8. From the results, it is found that multiple λ-type shock waves have no significant effect on the airfoil surface pressure distribution, the separated shear layer effect is dominant in the surface pressure change and aerodynamic characteristics.展开更多
The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish th...The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish the complete unsteadily aeroelastic coupling analytical model of the tiltrotor. The stability of the tiltrotor in the helicopter mode is analyzed aiming at a semi span soft-inplane tihrotor model with an elastic wing. Parametric effects of the lag stiffness of blades and the flight speed are analyzed. Numerical simulations demonstrate that the multibody analytical model can analyze the aeroelastic stability of the tiltrotor aircraft in the helicopter mode.展开更多
An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from ...An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°展开更多
The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carr...The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carrying one heavy load, an inertia coordinate system and body coordinate systems of each sub-system are established. A nonlinear force model is established too. The equilibrium computation results can be regarded as the reference control inputs of the flight control system under hovering or low-speed flight condition. After the establishment of a translation kinematics model and a posture kinematics model, a coupling dynamics model of the multiple helicopter system is set up. The results can also be regarded as the base to analyze stabilization and design a controller for a close-coupling multiple helicopters harmony operation system.展开更多
A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the...A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the model reference adaptive control law is designed and an adaptive compensator is used for improving its self-re- pairing capability. To enhance anti-interference capability of helicopter, quantum control feedforward is added be- tween fault and disturbance. Simulation results illustrate the effectiveness and feasibility of the approach.展开更多
Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to...Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to the structural fatigue load and the strength. The method of an optimal control law to suppress the gust disturbance for helicopters is presented. The optimization requires the minimization of the vertical overload at the pilot′s seat, the attitude variation and the control energy consumption under the gust disturbance. Based on the original control system, the new system can be easily realized by adding a vertical speed feedback passage. In order to develop the real-time operational flight control system, the optimized control law is written in C language. The hybrid simulations prove that the performance of gust response alleviation and the efficiency of digitalization are satisfactory.展开更多
Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to ext...Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to extract harmonic frequencies from really measured helicopter acoustic signal and an algorithm based on the SVD TLS was used. Results ESPRIT correctly extracted harmonic frequencies of helicopter using the data of limited length under the variousflight conditions. Conclusion ESPRIT is an effective method of extracting harmonic frequencies and using harmonic frequencies of helicopter acoustic signal to recognize helicopter is feasible.展开更多
Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical ...Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical model that has features such as rapidness, reliability and precision, because there is no unique and precise expression to some sophisticated phenomenon of helicopter. In this paper a fuzzy helicopter flight model is constructed based on the flight experimental data. The fuzzy model, which is identified by fuzzy inference, has characteristics of computed rapidness and high precision. In order to guarantee the precision of the identified fuzzy model, a new method is adopted to handle the conflict fuzzy rules. Additionally, using fuzzy clustering technology can effectively reduce the number of rules of fuzzy model, namely, the order of the fuzzy model. The simulation results indicate that the method of this paper is effective and feasible.展开更多
Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is pre...Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is presented. The model inversion is under the hover condition. And the adaptive control law based on the neural network is designed to guarantee the boundedness of tracking error and control signals. Simulation results demonstrate that the nonlinear neural network augmented model inversion can self-adapt to the uncertainty and modeling errors of unmanned helicopters. Results are compared while the parameters of PD controller and robustness items are changed.展开更多
The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled ...The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.展开更多
The circulation contr ol tail boom is a new technology used as an anti-torque system for the no tail rotor helicopter. In order to study the effect of the different parameters on th e circulation control tail boom, t...The circulation contr ol tail boom is a new technology used as an anti-torque system for the no tail rotor helicopter. In order to study the effect of the different parameters on th e circulation control tail boom, two models with the different circulation contr ol tail booms are designed and constructed. The experiments are introduced in th is paper by measuring the pressure distribution on the models in wind tunnel and under main rotor, respectively. By analyzing and discussing the experimental re sults, the effects of the momentum coefficient, the slot′s geometry parameters and the freestream velocity on the aerodynamic forces of the circulation control tail boom have been expounded, and the reasonable explanations about the test r esult have been given in theory.展开更多
The 2-dimensional equivalent model used in helicopter ground resonance analysis is convenient for its simplicity and clarity. The equivalent body mass, the stiffness and the damping are derived from the modal charaCte...The 2-dimensional equivalent model used in helicopter ground resonance analysis is convenient for its simplicity and clarity. The equivalent body mass, the stiffness and the damping are derived from the modal charaCteristics. In this paper, a 3-dimensional space model is constructed to analyse the ground resonance.There is little difference between the calculated results of the 2-dimensional equivalent model and those of the 3-dimensional space model. Hence, the 2-dimensional model is verified for the practical application. Generally, the linear lead-lag damping of the equivalent rotor blade at the modal frequency is derived from the lead-lag damper characteristics by a procedure of iteration. The transformation of the imaginary part of the complex modulus of the viscoelastic damper into a linear velocity damping is complicated in analysis. A method to use the complex modulus directly in the ground resonance analysis is derived and verified.展开更多
In this paper, in order to clarify the gains of the shroud in the shrouded tail rotor system, a thrust division factor q , which represents the ratio of the shroud thrust to the total thrust of the shrouded tail ro...In this paper, in order to clarify the gains of the shroud in the shrouded tail rotor system, a thrust division factor q , which represents the ratio of the shroud thrust to the total thrust of the shrouded tail rotor, is introduced. With the help of q , the slipstream theory for the static and axial flow states of the shrouded tail rotor are fully derived. Based on the sliptream theory, the variations of the thrust, power and disk area against q for different cases are emphatically analysed and the comparisons between a shrouded tail rotor and an isolated one are made. It is shown that, although the shroud can provide as much as 50% of the total thrust of shrouded tail rotor for the static state, the thrust gains of the shroud rapidly decrease for the axial flow state, which depends on the flow velocity ratio.展开更多
The feasibility of providing the tank a limited anti helicopter ability with gun launched missile is studied. A type of simulation model of gun launched missile against attack helicopters is established. The simula...The feasibility of providing the tank a limited anti helicopter ability with gun launched missile is studied. A type of simulation model of gun launched missile against attack helicopters is established. The simulation and the parameter optimization of missile control system under various circumstances are done. The gun launched missile can directly hit the helicopters in the typical tracks, all the missdistances are less than 1?m and the maximum overload is less than available overload. Gun launched missile is a feasible choice for tanks against attack helicopters.展开更多
Damping faults in a helicopter rotor hub are diagnosed by using vibration signals from the fuselage. Faults include the defective lag damper and raspings in its flap and feathering hinges. Experiments on the diagnosis...Damping faults in a helicopter rotor hub are diagnosed by using vibration signals from the fuselage. Faults include the defective lag damper and raspings in its flap and feathering hinges. Experiments on the diagnosis of three faults are carried out on a rotor test rig with the chosen fault each time. Fuselage vibration signals from specified locations are measured and analyzed by the fast Fourier transform in the frequency domain. It is demonstrated that fuselage vibration frequency spectra induced by three faults are different from each other. The probabilistic neural network （PNN） is adopted to detect three faults. Results show that it is feasible to diagnose three faults only using fuselage vibration data.展开更多
A practical survey on engineering implementation of flight control laws on helicopter engineering simulators is proposed.Advances of helicopter engineering simulators are introduced.Practical flight control technologi...A practical survey on engineering implementation of flight control laws on helicopter engineering simulators is proposed.Advances of helicopter engineering simulators are introduced.Practical flight control technologies are reviewed,with an emphasis on discussing the corresponding engineering simulation programs.Finally,the difficulties of implementing advanced control technologies are addressed,and the future development of helicopter engineering simulators are highlighted.展开更多
基金partially supported by the DEEPCOBOT project under Grant 306640/O70 funded by the Research Council of Norway.
文摘In this paper,we consider a 2-degrees-of-freedom(DOF)helicopter system subject to long input delays and uncertain system parameters.To address the challenges including unknown system parameters and input delays in control design,we develop an adaptive predictor-feedback control law to achieve trajectory tracking.Stability of the closed-loop system is further established,where the tracking errors are shown to converge towards zero.Through simulation and experiments on the helicopter system,we illustrate that tracking of a desired trajectory is achieved with the proposed control scheme.
基金supported in part by the National Natural Science Foundation of China(62273112,62061160371,61933001,51905115)the Science and Technology Planning Project of Guangzhou City(202201010758)+2 种基金the Guangzhou University-Hong Kong University of Science and Technology Joint Research Collaboration Fund(YH202205)the Open Research Fund from the Guangdong Laboratory of Artificial Intelligence and Digital Economy(Shenzhen(SZ))(GML-KF-22-27)the Korea Institute of Energy Technology Evaluation and Planning Through the Auspices of the Ministry of Trade,Industry and Energy,Republic of Korea(20213030020160)。
文摘Helicopter systems present numerous benefits over fixed-wing aircraft in several fields of application.Developing control schemes for improving the tracking accuracy of such systems is crucial.This paper proposes a neural-network(NN)-based adaptive finite-time control for a two-degree-of-freedom helicopter system.In particular,a radial basis function NN is adopted to solve uncertainty in the helicopter system.Furthermore,an event-triggering mechanism(ETM)with a switching threshold is proposed to alleviate the communication burden on the system.By proposing an adaptive parameter,a bounded estimation,and a smooth function approach,the effect of network measurement errors is effectively compensated for while simultaneously avoiding the Zeno phenomenon.Additionally,the developed adaptive finite-time control technique based on an NN guarantees finitetime convergence of the tracking error,thus enhancing the control accuracy of the system.In addition,the Lyapunov direct method demonstrates that the closed-loop system is semiglobally finite-time stable.Finally,simulation and experimental results show the effectiveness of the control strategy.
基金supported by Rotor Aerodynamics Key Laboratory[Grant No.RAL202102-4].
文摘The so-called coaxial compound helicopter features two rigid coaxial rotors,and possesses high-speed capabilities.Nevertheless,the small separation of the coaxial rotors causes severe aerodynamic interactions,which require careful analysis.In the present work,the aerodynamic interaction between the various helicopter components is investigated by means of a numerical method considering both hover and forward flight conditions.While a sliding mesh method is used to deal with the rotating coaxial rotors,the Reynolds-Averaged Navier-Stokes(RANS)equations are solved for the flow field.The Caradonna&Tung(CT)rotor and Harrington-2 coaxial rotor are considered to validate the numerical method.The results show that the aerodynamic interaction of the two rigid coaxial rotors significantly influences hover’s induced velocity and pressure distribution.In addition,the average thrust of an isolated coaxial rotor is smaller than that of the corresponding isolated single rotor.Compared with the isolated coaxial rotor,the existence of the fuselage results in an increment in the thrust of the rotors.Furthermore,these interactions between the components of the considered coaxial compound helicopter decay with an increase in the advance ratio.
基金supported in part by the National Natural ScienceFoundation of China (U2013201)the National Science Fund for Distinguished Young Scholars (61825302)the Postgraduate Research&Practice Innovation Program of Jiangsu Province (KYCX20_0208)。
文摘In this paper, a disturbance observer-based safe tracking control scheme is proposed for a medium-scale unmanned helicopter with rotor flapping dynamics in the presence of partial state constraints and unknown external disturbances. A safety protection algorithm is proposed to keep the constrained states within the given safe-set. A second-order disturbance observer technique is utilized to estimate the external disturbances. It is shown that the desired tracking performance of the controlled unmanned helicopter can be achieved with the application of the backstepping approach, dynamic surface control technique, and Lyapunov method. Finally, the availability of the proposed control scheme has been shown by simulation results.
文摘High-speed rotor rotation under the low-density condition creates a special low-Reynolds compressible flow around the rotor blade airfoil where the compressibility effect on the laminar separated shear layer occurs. However, the compressibility effect and shock wave generation associated with the increase in the Mach number (M) and the trend change due to their interference have not been clarified. The purpose is to clear the compressibility effect and its impact of shock wave generation on the flow field and aerodynamics. Therefore, we perform a two-dimensional unsteady calculation by Computational fluid dynamics (CFD) analysis using the CLF5605 airfoil used in the Mars helicopter Ingenuity, which succeeded in its first flight on Mars. The calculation conditions are set to the Reynolds number (Re) at 75% rotor span in hovering (Re = 15,400), and the Mach number was varied from incompressible (M = 0.2) to transonic (M = 1.2). The compressible fluid dynamics solver FaSTAR developed by the Japan aerospace exploration agency (JAXA) is used, and calculations are performed under multiple conditions in which the Mach number and angle of attack (α) are swept. The results show that a flow field is similar to that in the Earth’s atmosphere above M = 1.0, such as bow shock at the leading edge, whereas multiple λ-type shock waves are observed over the separated shear layer above α = 3° at M = 0.80. However, no significant difference is found in the C<sub>p</sub> distribution around the airfoil between M = 0.6 and M = 0.8. From the results, it is found that multiple λ-type shock waves have no significant effect on the airfoil surface pressure distribution, the separated shear layer effect is dominant in the surface pressure change and aerodynamic characteristics.
文摘The muhi-body analysis of the aeroelastic stability of the tiltrotor aircraft is presented. Muhi-body dynamic differential equations are combined with the equations of the unsteady dynamic inflow model to establish the complete unsteadily aeroelastic coupling analytical model of the tiltrotor. The stability of the tiltrotor in the helicopter mode is analyzed aiming at a semi span soft-inplane tihrotor model with an elastic wing. Parametric effects of the lag stiffness of blades and the flight speed are analyzed. Numerical simulations demonstrate that the multibody analytical model can analyze the aeroelastic stability of the tiltrotor aircraft in the helicopter mode.
文摘An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°
基金National Natural Science Foundation of China(60475039)
文摘The particularity and practicality of harmony operations of close-coupling multiple helicopters indicate that the researches on it are urgent and necessary, Using the model that describes two hovering helicopters carrying one heavy load, an inertia coordinate system and body coordinate systems of each sub-system are established. A nonlinear force model is established too. The equilibrium computation results can be regarded as the reference control inputs of the flight control system under hovering or low-speed flight condition. After the establishment of a translation kinematics model and a posture kinematics model, a coupling dynamics model of the multiple helicopter system is set up. The results can also be regarded as the base to analyze stabilization and design a controller for a close-coupling multiple helicopters harmony operation system.
基金Supported by the National Natural Science Foundation of China(61074080)the Innovation Foundation for Aeronautical Science and Technology(08C52001)~~
文摘A direct self-repairing control approach is proposed for helicopter via quantum control techniques and adaptive compensator when some complex faults occur. For a linear varying-parameter helicopter control system, the model reference adaptive control law is designed and an adaptive compensator is used for improving its self-re- pairing capability. To enhance anti-interference capability of helicopter, quantum control feedforward is added be- tween fault and disturbance. Simulation results illustrate the effectiveness and feasibility of the approach.
文摘Gust response alleviation is very important for helicopters which have strong coupling and vibration. Gust disturbance not only influences the ride quality and the precision of the weapon delivery, but also affects to the structural fatigue load and the strength. The method of an optimal control law to suppress the gust disturbance for helicopters is presented. The optimization requires the minimization of the vertical overload at the pilot′s seat, the attitude variation and the control energy consumption under the gust disturbance. Based on the original control system, the new system can be easily realized by adding a vertical speed feedback passage. In order to develop the real-time operational flight control system, the optimized control law is written in C language. The hybrid simulations prove that the performance of gust response alleviation and the efficiency of digitalization are satisfactory.
文摘Aim To extract harmonic frequencies of helicopter acoustic signal as features for hel icopter identification. Methods Estimation of signal parameters via rotational invariance techniques(ESPRIT) was selected to extract harmonic frequencies from really measured helicopter acoustic signal and an algorithm based on the SVD TLS was used. Results ESPRIT correctly extracted harmonic frequencies of helicopter using the data of limited length under the variousflight conditions. Conclusion ESPRIT is an effective method of extracting harmonic frequencies and using harmonic frequencies of helicopter acoustic signal to recognize helicopter is feasible.
文摘Helicopter mathematical model mainly depends on design helicopter control system, flight simulator, and real time control simulation system. But it is difficult to establish a helicopter flight dynamics mathematical model that has features such as rapidness, reliability and precision, because there is no unique and precise expression to some sophisticated phenomenon of helicopter. In this paper a fuzzy helicopter flight model is constructed based on the flight experimental data. The fuzzy model, which is identified by fuzzy inference, has characteristics of computed rapidness and high precision. In order to guarantee the precision of the identified fuzzy model, a new method is adopted to handle the conflict fuzzy rules. Additionally, using fuzzy clustering technology can effectively reduce the number of rules of fuzzy model, namely, the order of the fuzzy model. The simulation results indicate that the method of this paper is effective and feasible.
文摘Adaptive flight control technology, feedback linearization, model inversion theory are reviewed and the error dynamic characteristics are analyzed, and an adaptive on-line neural network attitude control system is presented. The model inversion is under the hover condition. And the adaptive control law based on the neural network is designed to guarantee the boundedness of tracking error and control signals. Simulation results demonstrate that the nonlinear neural network augmented model inversion can self-adapt to the uncertainty and modeling errors of unmanned helicopters. Results are compared while the parameters of PD controller and robustness items are changed.
文摘The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.
文摘The circulation contr ol tail boom is a new technology used as an anti-torque system for the no tail rotor helicopter. In order to study the effect of the different parameters on th e circulation control tail boom, two models with the different circulation contr ol tail booms are designed and constructed. The experiments are introduced in th is paper by measuring the pressure distribution on the models in wind tunnel and under main rotor, respectively. By analyzing and discussing the experimental re sults, the effects of the momentum coefficient, the slot′s geometry parameters and the freestream velocity on the aerodynamic forces of the circulation control tail boom have been expounded, and the reasonable explanations about the test r esult have been given in theory.
文摘The 2-dimensional equivalent model used in helicopter ground resonance analysis is convenient for its simplicity and clarity. The equivalent body mass, the stiffness and the damping are derived from the modal charaCteristics. In this paper, a 3-dimensional space model is constructed to analyse the ground resonance.There is little difference between the calculated results of the 2-dimensional equivalent model and those of the 3-dimensional space model. Hence, the 2-dimensional model is verified for the practical application. Generally, the linear lead-lag damping of the equivalent rotor blade at the modal frequency is derived from the lead-lag damper characteristics by a procedure of iteration. The transformation of the imaginary part of the complex modulus of the viscoelastic damper into a linear velocity damping is complicated in analysis. A method to use the complex modulus directly in the ground resonance analysis is derived and verified.
文摘In this paper, in order to clarify the gains of the shroud in the shrouded tail rotor system, a thrust division factor q , which represents the ratio of the shroud thrust to the total thrust of the shrouded tail rotor, is introduced. With the help of q , the slipstream theory for the static and axial flow states of the shrouded tail rotor are fully derived. Based on the sliptream theory, the variations of the thrust, power and disk area against q for different cases are emphatically analysed and the comparisons between a shrouded tail rotor and an isolated one are made. It is shown that, although the shroud can provide as much as 50% of the total thrust of shrouded tail rotor for the static state, the thrust gains of the shroud rapidly decrease for the axial flow state, which depends on the flow velocity ratio.
文摘The feasibility of providing the tank a limited anti helicopter ability with gun launched missile is studied. A type of simulation model of gun launched missile against attack helicopters is established. The simulation and the parameter optimization of missile control system under various circumstances are done. The gun launched missile can directly hit the helicopters in the typical tracks, all the missdistances are less than 1?m and the maximum overload is less than available overload. Gun launched missile is a feasible choice for tanks against attack helicopters.
文摘Damping faults in a helicopter rotor hub are diagnosed by using vibration signals from the fuselage. Faults include the defective lag damper and raspings in its flap and feathering hinges. Experiments on the diagnosis of three faults are carried out on a rotor test rig with the chosen fault each time. Fuselage vibration signals from specified locations are measured and analyzed by the fast Fourier transform in the frequency domain. It is demonstrated that fuselage vibration frequency spectra induced by three faults are different from each other. The probabilistic neural network （PNN） is adopted to detect three faults. Results show that it is feasible to diagnose three faults only using fuselage vibration data.
基金supported by the Fundamental Research Funds for the Central Universities (No. XBC16010)
文摘A practical survey on engineering implementation of flight control laws on helicopter engineering simulators is proposed.Advances of helicopter engineering simulators are introduced.Practical flight control technologies are reviewed,with an emphasis on discussing the corresponding engineering simulation programs.Finally,the difficulties of implementing advanced control technologies are addressed,and the future development of helicopter engineering simulators are highlighted.
